Large gas turbine engines are widely used for aircraft propulsion and for ground based power generation. Such large gas turbine engines are of the axial type, and include a compressor section, a combustor section, and a turbine section, with the compressor section normally preceded by a fan section. Each of the fan, compressor, and turbine sections comprises a plurality of disks mounted on a shaft, with a plurality of airfoil shaped blades projecting radially therefrom. A hollow case surrounds the various engine sections. Located between the disks and projecting inward from the case assembly which surrounds the disks are a plurality of stationary vanes. During operation of the fan, compressor, and turbine sections, axially flowing gases alternately contact moving blades and the stationary vanes. In the fan and compressor sections, air is compressed and the compressed air is combined with fuel and burned in the combustion section to provide high pressure, high temperature gases which flow through the turbine section, where energy is extracted by causing the bladed turbine disks to rotate. A portion of this energy is used to operate the compressor section and the fan section.
Engine efficiency depends to a significant extent upon minimizing leakage by control of the gas flow in such a manner as to maximize interaction between the gas stream and the moving and stationary airfoils. A major source of inefficiency is leakage of gas around the tips of the compressor blades, between the blade tips and, the engine case. Accordingly, means to improve efficiency by reduction of leakage are increasingly important. Although a close tolerance fit may be obtained by fabricating the mating parts to a very close tolerance range, this fabrication process is extremely costly and time consuming. Further, when the mated assembly is exposed to a high temperature environment and high stress, as when in use, the coefficients of expansion of the mating parts may differ, thus causing the clearance space to either increase or decrease. The latter condition would result in a frictional contact between blades and housing, causing elevation of temperatures and possible damage to one or both members. On the other hand, increased clearance space would permit gas to escape between the compressor blade and housing, thus decreasing efficiency.
One means to increase efficiency is to apply a coating of suitable material to the interior surface of the compressor housing, to reduce leakage between the blade tips and the housing. Various coating. techniques have been employed to coat the inside diameter of the compressor housing with an abradable coating that can be worn away by the frictional contact of the compressor blade, to provide a close fitting channel in which the blade tip may travel. Thus, when subjecting the coated assembly to a high temperature and stress environment, the blade and the case may expand or contract without resulting in significant gas leakage between the blade tip and the housing. This abradable coating technique has been employed to not only increase the efficiency of the compressor, but to also provide a relatively speedy and inexpensive method for restoring excessively worn turbine engine parts to service.
As generally mentioned in U.S. Pat. Nos. 3,879,831 to Rigney et al, and 3,084,064 to Cowden et al, abradable seals must have a peculiar combination of properties. They must be resistant to erosion from the high velocity, high temperature gas streams which at times may carry fine particulate matter with them. However, they must also be subject to removal (i.e. abrading) when contacted by the tip of a high speed blade, so that the tip of the blade is not degraded. It is critical that the housing coating abrade rather than wear the blade tip, since a decrease in blade tip size will increase clearance between the blade tip and the housing all around the circumference, resulting in a greater increase in gas leakage than would result from abrasion of only a small arc of the coating around the circumference of the housing. Conventionally, the tip of the blade is coated with a highly erosion resistant material.
The abradable coating must also be structurally sound to resist failure at points other than where contacted by the blade tip, must resist the thermal and vibratory strains imposed upon it in use, and must be readily fabricated in a reproducible and cost efficient manner. Considerable effort has gone into the development of abradable seals having the desired combination of properties. The present invention is reflective of that continuing effort.
One form of abradable seal developed in the past was a porous structure, obtained by use of a fugitive material in the precursor article. In the prior art, pressing and sintering and other metallurgical techniques have been used together with thermal spraying to produce porous structures. Metal deposits with densities as low as 75-85 percent may be applied by plasma spraying. However, to obtain densities lower than this, which were formerly believed to be desirable for abradable seals, it was necessary to incorporate non-metallic materials. Most preferably, a fugitive material such as a water soluble salt or a heat-decomposable polymer was sprayed with the metal, and then subsequently removed. For example, an abradable seal structure is prepared in accordance with the teachings of Otfinoski et al, U.S. Pat. No. 4,664,973, who teaches spraying a polymethylmethacrylate resin and a nichrome metal, and then removing the resin by heating the resultant structure to a temperature of about 315.degree. C.
Another form of abradable seal is that prepared by the teachings of Rigney et al, U.S. Pat. No. 3,879,831. This patent discloses an abradable material having a composition of 60-80 percent nickel, 2-12 percent chromium, 1-10 percent cobalt, 4-20 percent aluminum, and 3-15 percent inert material such as diatomaceous earth, boron nitride, silica glass, mica, etc. Up to 3 percent of a metal such as yttrium, hafnium, or lanthanum may also be present. The abradable materials produced by this reference are characterized by a high degree of porosity, oxidation resistance, low thermal conductivity, and the ability to be abraded away cleanly in a localized area.
Similarly, U.S. Pat. No. 3,084,064 deals with the preparation of abradable coatings on turbine surfaces by flame spraying nichrome and from 2 to 20 weight percent of a finely divided powder of a high melting material such as boron nitride, carbon, graphite, or magnesium oxide. The abradable characteristics of this coating are believed to be due to the dispersed material preventing formation of a solid, dense, strongly cohesive metal phase. In other words, the high melting powder permits the surface to easily flake off in relatively uniform particles when subjected to an abrading force.
Although these various methods produce abradable coatings usable for turbine applications and the like, they have disadvantages of either providing coatings which are hard to chip off in small discreet amounts by friction contact so as to provide a well defined blade tip channel having no large cavities through which gases may escape, or producing an interconnected porous surface layer which in itself permits the escape of gases, thus lowering efficiency.
Accordingly, it is an objective of the present invention to provide an improved seal system which contributes to engine efficiency by providing a compressor seal, which while abradable and smooth, is impermeable to gas flow. It is a further object of the present invention to provide a compressor seal comprising plasma sprayed metal matrix, a lubricating amount of boron nitride second phase, and porosity which is not interconnected.